This invention relates generally to gas turbine engines and more particularly to monitoring the thermal state in such engines.
Gas turbine engines are used for a wide variety of aeronautical, marine and industrial applications. Generally, a gas turbine engine includes a compressor that provides pressurized air to a combustor, wherein the air is mixed with fuel and the mixture is ignited for generating hot combustion gases. These gases flow downstream to a turbine section that extracts energy therefrom to drive the compressor and provide useful work. In many applications, gas turbine engines are routinely subject to various maintenance procedures as part of their normal operation. To aid in the provision of such maintenance services, monitoring systems are often employed to provide diagnostic monitoring of the gas turbine engine. These systems commonly include performance monitoring equipment that collects relevant trend and fault data used for diagnostic trending. In diagnostic trend analysis, certain process data (such as exhaust gas temperature, fuel flow, rotor speed and the like) that are indicative of overall engine performance and/or condition are compared to a parametric baseline for the gas turbine engine. Any divergence of the raw trend data from the parametric baseline may be indicative of a present or future condition that requires maintenance.
For example, modem aircraft currently operated by commercial airlines typically employ an onboard data acquisition system for collecting digital flight data to use in diagnostic monitoring. In such systems, a number of sensors distributed throughout the aircraft and engines provide data signals representative of the performance of the aircraft and its engines. Such data can be recorded onboard and accessed later by ground maintenance personnel or, alternatively, can be remotely transmitted to ground locations during flight operations for real-time processing.
Engine condition monitoring techniques typically use a screening process to identify various phases of operation and then extract specific data during the flight phases of interest. Currently, data collection is conducted during flight phases such as take off, climb and steady cruise, because these are the phases during which engine anomalies are most likely to be detected. Data collected during the takeoff phase can be strongly influenced by the engine""s thermal state at engine start-up. For example, bearing and rotor clearances are generally more open during a cold rotor start (e.g., the first start of the day) than during a hot start (e.g., a start after a recently concluded flight). This means that rubbing and rotor bow are more likely to occur during hot starts.
Accordingly, it is desirable to be able to monitor engine thermal state characteristics in gas turbine engines prior to engine start-up for the purpose of normalizing general engine performance characteristics.
The above-mentioned need is met by the present invention, which provides a method and system for monitoring engine performance in a gas turbine engine in which a plurality of sensors are used to sense data related to the operation and performance of the gas turbine engine. Selected data parameters from the sensed data are continuously sampled prior to completion of an engine start sequence. The selected data parameters are then evaluated to determine whether specific criteria have been met, and evaluation data are captured whenever the criteria are met. The evaluation data are used to normalize engine performance data to a particular thermal state of the engine. The normalized engine performance data are then trended by comparing it to a parametric baseline for the gas turbine engine.
The present invention and its advantages over the prior art will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings.